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Are there regulations for unmanned aerial systems avionics?
I know that there are documents such as DO-178B, DO-178C and DO-254 that provide guidance and requirements for the development of avionics in general.
In my opinion it's hard to believe that avionics developed for unmanned aerial systems (left) are subject to the same technical requirements than those on board manned aircraft (right).

Therefore I ask:

Are there any specific regulations for avionics on board unmanned aerial systems?
If so, how do they differ from those specified in documents such as DO-178B, DO-178C and DO-254?

I doubt DO-178 or DO-254 has been applied to any consumer grade UAS (not sure about military). DO-178 and DO-254 are meant for safety critical software / electronic hardware. As UAS so far haven't been deemed to require safety critical software, I expect that the software developed for UAS are done to each company's processes/standards

If there aren't yet, there will be once a falling, out of control camera drone kills or injures a person or property...

The flight control software on military grade UAV/UAS has DO-178/DO-254 requirements, per briefings I attended, and also that software provided for embedding in those had to meet those and other requirements.
I am not an expert on the UAS platform we used (Global Hawk RQ-4) but it is safe to say that it is a complex bird with multiple capabilities. A typical mission flight crew consisted of a launch and recovery pilot, and mission pilots and sensor operators, in sufficient numbers to handle extended missions.
I am unaware of any consumer grade UAS utilizing those standards, and doubt there are many, if any at all, due to the increased development cost. I am unaware of any type certificated sUAS.
To give you an example, I had a certain nav sensor related function integrated for VFR use on a sensor platform (Twin Otter).  The development cost was about 4.5 million. If we were to apply the standards so that we could operate that platform in an IFR environment, the estimated cost was about 9.5 to 11 million.  Sure there were just a few additional capabilities, but a big part of it was systems engineering and test.

Thanks a lot for your answer. I wonder under which circumstances such requirements need to be met and when not, specially for civil applications. Do you happen to know?

The answer is beyond what can readily be covered here.  However, when dealing with systems which are large (type certificated), that act autonomously, that go beyond visual line of sight, that integrate with airspace at altitudes where manned aircraft normally operate, are among the criteria which influence the need for reliable software. If you want to dig further, I would look at the type certification. I would expect that all the military employed and internationally dispatched UAS require avionics grade software development.


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What determines spanwise placement of spoilers for roll control?
Specific application of spoilers was discussed in What is the benefit of spoilerons compared to ailerons? As I remember it, 40yrs ago gliders had spoilers close to the wing tips - perhaps slightly inboard of the normal aileron position.
Current aircraft design places them mid-wing or even next to the wing root. What design principles favor a particular span-wise location.
To be specific; I am interested in roll control not drag.

But are those spoilers near wing root actually being used for roll control? I don't think so. The one on your picture certainly isn't—the spoilers are deployed symmetrically. And I don't think they use them that way at all. Airliners do, but that use always deploys the outboard panels only.

It has to do with the aerodynamic and the coefficient of the drag and lift. When the spoilers are close to the wingtips (end of the wing) the coefficient of the drag is less than when they are at the middle of the wing. The same happens when they are at the beginning of the wing. The drag coefficient is less. So at the middle of the wing, the drag coefficient is the highest, which fulfills the role of the spoilers. This is most likely the reason why people changed the position from the end of the wing to the middle, as this apparently is better.

This is most likely because the length of a glider wing is smaller compared to an airliner's, and because of this, the position of the spoilers don't matter as much.

Thanks for the comment, I saw that I had left ambiguity as to the specific purpose of the spoilers I was interested in - I am specifically interested in roll control and edited my question to reflect this (my fault not yours :). I certainly understand placement for DRAG only purposes. But mid-wing or inboard seems to defeat effective roll control.

On airliners, spoilers serve three purposes, roll enhancement (assist the ailerons when greater roll rate is commanded), speed brakes (help slow down and descend faster), and ground spoilers (kill lift in order to transfer weight from the wings to the wheels for more effective braking). Spoiler span-wise location is a compromise to accomplish all three goals. Also, the outboard spoilers are primary for roll control and the most inboard spoilers may only be used on the ground. Spoilers are not placed in front of the ailerons because they would interfere with roll control. The chord-wise location leaves the wing box clear for fuel tanks.

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Why aren't Rogallo wings or parafoils used in aircraft with fuselages?
I was wondering why no aircraft with a cockpit and fuselage use Rogallo wings or parafoils. I'm guessing they would have the advantage of low weight, collapsibility (storage), and low cost.

Larger aircraft: Metallic wings
How much fuel could you store in the Rogallo wings if you need some?

Source
The wings provide lift, but not only. In large aircraft they usually receive thrust from the engines, in order to transfer it to the fuselage, they are also fitted with hyper-sustentation devices and aerobrakes. And finally they can transmit a good part of the aircraft weight to the outer landing gears when the aircraft is on the ground.

Source
The latter functions are not compatible with a flexible structure. Only metallic or composite wings are used. They can be adjusted for an accurate envelope with optimized performances.

Smaller aircraft: Wood and canvas or metallic wings
The structure can be stiffened, this is commonly done, and leads to the concept of wood and canvas, as the DR-400 and many other GA training aircraft (all first aircraft were built this way):

Source
However these airframes don't scale well, and are not compatible with fast GA aircraft, leaving alone airliners speed range. Larger and faster aircraft, which anyway must protect their wing tanks, use metallic wings.

Ultralights and paragliders: Only categories with Rogallo wings
The advantages provided by flexible wings (easy storage, light weight, low cost) is a significant choice criteria for use in ultralights and paragliders, and of course for parachutes. These advantages overcome their limited aerodynamic performances, which is of lesser consideration for vehicles that are not specially aerodynamically designed.

Source

To summarize: The flexible wings are more appropriate for smaller and slower aircraft, in particular those not requiring wing tanks, wing control surfaces, or wing landing gears. Namely ultralights.

It might be useful to mention that airliners require high-lift devices on the wings to have reasonable take-off & landing speeds.  A Rogallo wing cannot be deformed in such a manner...

The wings in the aircraft fulfill a number of functions, which would be quite difficult and inefficient with the Rogallo wings (or parafoils, for that matter), like:

L/D ratio- The main purpose of wings is for efficient generation of lift i.e. with minimum drag. The L/D ratio of modern airliners are anywhere between 15-20, while in case of the parafoil/Rogallo wing the values are in the range of 5. One thing to note is that an important way to improve the L/D ratio in normal wings is to increase the aspect ratio, which is severely restricted in case of these flexible wings- even for aircraft design, flutter is a serious problem.

L/D ratio of flexible wings, table from Gliding Parachutes for Land Recovery of Space Vehicle by W. H. Eilertson

Fuel carrying capacity- Apart from creating lift, an important function of aircraft wings is to carry fuel, which isn't practical in case of flexible wings- this alone would be enough to restrict these devices to smaller aircraft.

Controls- The wings house the controls, which would not be possible in case of flexible wings (future combat aircraft are expected to have wing morphing, but is still a bit far away). The control of flexible wings are by using cables, which are suitable only for small aircraft. Another issue is redundancy, which would be quite difficult with these exposed cables.

Control of rogallo wing using cables; from NASA's Gemini program; image from vintagespace.files.wordpress.com

Engine mounting- Wings serve as an important place for mounting engines, which would not be possible in this case- the aircraft are restricted to single or inline engines.

Landing gear mounting/retraction- in some cases, the landing gear is mounted/retracted into the wing, which again would not be possible in case of flexible wings.

Safety- The flexible wings do not offer the same level of safety as the rigid (elastic) ones. What if they fail to deploy properly or start to retract in flight? NASA carried out research on the usage of flexible wings for the spacecraft landing programs which brought out a number of issues related to structural safety and deployment:

The wing’s structural problems persisted. Further drop tests saw more sails fall apart ...
Wing deployment tests were also proving problematic. The sail wasn’t deploying consistently, at times not opening fully and failing to inflate with enough time...

Even though most of these problems were fixed, it would have been difficult to convince regulatory authorities and paying public for their use.
It should be noted that most of the advantages of the flexible wings are precisely because of their limitations- it is collapsible for reduced storage and weigh less because it doesn't carry fuel or have control surfaces and doesn't have mountings for various things (like navigation lights etc).
It should noted that the people involved were quite conscious of the restrictions of this type of wing- For example, Francis M. Rogallo, the inventor of rogallo wing was quite cautious when he noted,

... if  we could discover how to make flexible wings that could be packaged and deployed somewhat like a parachute, such wings would have many new applications as well as replacing some parachutes and rigid wings.

if we could discover how to make flexible wings that could be packaged and deployed somewhat like a parachute -- in effect that's what we have today. It's just not Rogallo wings but parafoils instead. Turns out that Rogallo wings have really bad L/D ratio and are really inefficient. Plank-style parafoil wings are good enough for people to glide accross the Himalayas and consistently break sailing speed records.

@slebetman Note that parafoil was invented (atleast the patent was filed) after this comment was made.

Yep, that's why I said "in effect that's what we have today" to give a sense that one of the things he wished came true only not in the way he envisioned. (note: I had to read your comment several times to realise you didn't mean your comment which would mean that the parafoil was patented less than 2 hours ago)

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Is it legal to listen to aviation frequencies in the UK?
Am i allowed to purchase an ATC scanner and listen tho what is being said in the control tower? Thank you

Yes, you can purchase a scanner - these are freely available in the UK.
No, legally, you can't listen to air traffic control. However, in real life, this is generally not enforced. Go to any airshow or airport observation area and you'll find people using scanners openly.

Why is it illegal to listen to ATC in the U.K.?

@CarloFelicione It's part of a broad brush of regulations covering snail mail, electronic comms and radio (amongst other types) which make it illegal to receive communications not intended for you.


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How does supercruise thrust compare with static thrust ASL?
As I understand it, turbojet thrust decreases both with altitude due to rarefied air reducing massflow, and airspeed due to a smaller speed increase between intake and exhaust relative to the airframe. However aircraft designed for sustained supercruise such as Concorde or Blackbird had intakes that used various amounts of ram-effect pressure recovery, so I was wondering how much that would offset the effects of being fast and high i.e. what was Concorde's mach 2 max thrust compared compared to sea level dry thrust, and/or the Blackbird's mach 3.2 turboramjet thrust compared to take-off wet thrust?

Related: In what way are the Concorde's engines considered efficient?

Thanks for the link that's a pretty interesting read, but I was asking about thrust not efficiency.

According to the wiki page of the Rolls Royce/Snecma Olympus for the Concorde:

During cruise at Mach 2.02 each Olympus 593 was producing around 10,000 lbf of thrust... Eames (SAE Transactions 1991) however mentions the cruise thrust of each engine to be 6790 lbf.... The 10,000 lbf is perhaps the maximum thrust available at cruising speed (used during acceleration and climb just before transitioning to cruise).

TO thrust is for the 593-610-14-28 - Final version fitted to production Concorde, is listed as 32,000 lbf (142 kN) dry / 38,050 lbf (169 kN) reheat.

As I understand it, turbojet thrust decreases both with altitude due to rarefied air reducing massflow, and airspeed due to a smaller speed increase between intake and exhaust relative to the airframe.

Yes indeed - however, available thrust increases with airspeed, that is why ramjets work. And drag reduces with altitude, so the required thrust decreases.

The contributions to thrust of the parts of a turbojet installation for supersonic propulsion have been computed and presented in prof. Wittenberg's TUDelft D-32B lecture handout, Fig 3.9 (above). It can be seen that the contribution of the inlet rises exponentially with Mach number M, while that of the engine itself reduces above M = 1.5. Above M = 2 the aeroplane drags the engine along! Reason why the turbine was completely bypassed at the M=3 for the SR-71.

In the same lecture handout, the pic above shows the subdivision over the installation parts of thrust, from "Power for supersonic flight", Journal of the Royal Aeronautical Society, July 1964.

For Concorde, takeoff thrust:weight was 1:0.373. At cruise, the Concorde had a lift:drag ratio of 7:1. At cruise, thrust = drag. So cruise thrust is 1/7 = 0.1429
So, cruise:ASL is 0.14(rounded down for climb fuel burn):0.373 = around  1:3.8
For the SR-71, takeoff thrust:weight is 0.44:1. Using the Kuchemann supersonic L/D formula, we get
4(Mach(=3)+3)/Mach(=3)
=1:8
Now this looks too high. And running this formula for Concorde gave 10:1. So, having no better option (yet), I simply took 1:6 as the value. This would give us a cruise thrust:weight of around 1/6:1 = 0.167:1
0.167/0.44 = around 1:3.8 (again?)
The figures for the SR-71 may well be wrong.


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How would a Flying Fortress behave if it lost its empennage mid-flight?
The 1990's movie "Memphis Belle" depicts a Flying Fortress having the whole tail cut by a fighter impact on the rear part of fuselage, as a razor or chainsaw will do, the bomber airplane enters a flat descent with all 4 engines operating.
Is this type of accident possible? How would such a bomber behave with the empennage cut? Is there any way or chance to survive such an accident?

Depends on the extent of the damage as well as what systems are affected.  Lose the vertical fin, it's possible the aircraft could make it home but with limited yaw stability.  Lose the vertical fin and the tailplane, the airplane would nose over and depart from controlled flight due to longitudinal instability.

Any type of accident is possible. A Flying Fortress needs its tail in order to be controllable, steady flight is not possible without horizontal surfaces. Without vertical tail maybe, by differential engine thrust.

A video exists in YouTube about a B-52 losing the vertical fin of tail, and how they managed to survive. A Japanese 747 wasn't that lucky, all on board died.

Yeah and a DC-10 managed to use differential engine thrust for pitch control. Cannot be done on a Flying Fortress, the thrust vector is very much aligned with the CoG.


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What are those hatches on the engines of a B-2 that are open during takeoff?
I just watched this drool-inducing video of Northrop Grumman B-2 Spirit, And noticed that there are a couple of hatches near each engine, which were open:

I initially thought they were open for some sort of maintenance. But the video features a takeoff roll of the bomber (around 5:00 minute mark), and I again noticed that the hatches were open during the roll. However, they were retracted immediately after the aircraft left the ground (around 5:25 mark):

This makes me wonder if these hatches serve some special purpose. What are they, and why are they open during the takeoff roll?
Edit: Please don't say you don't see an aircraft, because the landing gear doors are down and radar signature is immense due to that. :)

Once I toured AMARC. They had displays of most of their aircraft with signs plainly identifying them.  This one sign said "Stealth bomber". It appeared to be an empty spot.

The refueling door rotating to close at the end is pretty neat - I didn't know it did that.

These are secondary intake doors which allow more air to reach the engines. They are opened by the engine control computer when the main intakes are not providing enough air for the power setting of the engine. In order to minimize radar cross section, the engines of the B-2 are buried in the wings and the intake needs to be curved. Also, and for the same reason, the intake lips need to be sharp-edged. Both factors increase the drag of the intake flow and reduce the mass flow of air to the engines. Like on other low-RCS aircraft, secondary intake openings are needed to increase this mass flow.
In automatic mode they will close when the aircraft accelerates above Mach 0.4. Below 43F, they remain closed for taxi and are only opened for take-off. A take-off with closed doors is possible according to this thread, but needs a reduced power setting to 80% for the engines below 50 kt speed.

Sounds like a similar function to "derumble doors", the suck-in panels on normal jet aircraft.

@JohnK Right, and that is why I linked to the JT3D-related answer in the first link.

Did the 727, Tu-154, L1011 and other S-duct aircraft have anything similar?

@Harper: No, but they have a much more favorable geometry: Round (almost in case of the 154) and with well rounded intake lips.

Doors for secondary air intakes. Buried engines need additional air during takeoff and low speed flight.

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How does the hinge mechanism of plug-type pressure doors work?
Many pressurized aircraft have the type of plug doors that open by first swinging inside,  rotating about at least two axes, then moving through the opening and ending up outside parallel with the fuselage.
I'm intrigued by the mechanism that makes this possible, but I have never been able to really see how these intricate hinges work. As a passenger it's hard to even witness the actual opening or closing of the door. Some drawings are to be found on the web, but none that I know of give a clear understanding of its workings. Does anybody know of a video or an animation that gives some insight?

This video has a clear shot of how a classic plug door (in this case of a B737) works (or search for flight attendant training videos).
You can see the door pops in (tipping a bit downwards) and then rotates on a axis in the center of the opening and then moves outside sideways. This is possible because of how the geometry works out, moving sideways negates the boor being wider than the opening and tipping down lets it fit height-wise.

Which means that the door is wider than the opening, but not higher.

@JanHudec the door tipping inward lets it be slightly higher (though a separate flap is used for the bottom and top see http://youtu.be/BVxPBeu2PaI?t=2m14s)

Why are manhole covers round? Because the holes are round. Actually, it's because a round cover can't fall through a round hole.

@rbp That must be at least one of the reasons why aircraft doors are not round.

The door you mentioned is called a Plug Door.

On some aircraft the plug door opens partially inward, and through a
  complex hinge design can be tilted to fit through the fuselage
  opening, or the door may have locking hinged panels at the top and
  bottom edges that can make it smaller than the opening, so that it may
  be swung outward.

This video shows opening and closing of such door from both outside and inside.

I know about those hinged panels, but afaik they are not used on all these plug doors. To me they seem to be a less effective solution, for the panels should open outward and consequently cannot (like the door itself) withstand cabin pressure without seperate latches.

A Google search of Plug Door Patents will yield many variants.  The ones you want are:

Outward-opening plug type door for aircraft
Double-decked aircraft with plug door
Folding, plug type aircraft door

I am going to build one of them as the entrance to my workshop - but I need to get the house built first.

I also liked this one: http://www.google.com/patents/US4199120
I could not put more than two links in my previous posted answer.

Providing more information from the links in your answer would be helpful.

Good luck Farhan! Looking forward to some pics or a video as soon as you're done. Good to build the house before making the plug door. That's what Boeing and Airbus found out to be the preferred order.


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Why didn't early jetliners use turbofans?
For subsonic operation, high-bypass turbofans seem to have essentially only advantages for efficiency, noise and takeoff thrust, at least until they reach the size where ground clearance or LP/fan speed mismatch become major issues. Yet early jetliners first used turbojets then only very low-bypass turbofans, despite the existence of turboprops meaning that using gas turbines for indirect propulsion was evidently understood.
Did they not fully realise the potential benefits at the time/thought they wouldn't work well, or was it just very difficult to build a jet with useful excess turbine power (to drive the fan) at the time?

Because a turbofan requires excess energy, and that means a high temperature core, and the temperature capability of early turbine blade material and combustion sections was insufficient. That's the reason why it was difficult to build an early gas turbine with excess power.  And it is still a limiting reason stopping higher bypass ratios, which increase propulsion efficiency.

Difficult when taking off excess power at the shaft, or also when taken from a power turbine?

In addition to the Germans appreciating the benefits of turbofans by WW2 (see Peter Kampf below), so did the British.  Frank Whittle understood by 1936 that it is more efficient to give a small acceleration to a large mass of air than a large acceleration to a small mass of air.  See this Stack Exch answer https://aviation.stackexchange.com/questions/35648/how-was-the-high-bypass-concept-invented?noredirect=1&lq=1

@rackandboneman.  Both, when taking power at the shaft, or from a power turbine. In either case, power is needed, and the thermodynamics is the same; both situations require a core that can accommodate the temperature required.

Production turbofans just weren't available. Once they came on the market and were mass-produced, jetliners switched to them.
While turboprops did exist at the time, they were (and are) a turbine engine driving a propeller in place of a piston engine, a straightforward combo of two techs. Turbojet engines were fairly complex machines for the era on their own.
A turbofan involves a large diameter fan, similar in design to an axial compressor, which is more difficult to build than simpler centrifugal compressors used in the first jets and most turboprops. It spins at higher rpm than a prop, and needs to be built and balanced to jet engine tolerances, with the fan tips very close to the casing, but not quite touching it.
The fan also needs to be driven by a core that has enough net power left to produce both thrust and torque, and to do so at low enough rpm to keep the fan (mostly) subsonic. This combination only brings considerable efficiency gains with a two-spool core, using a low-speed shaft to match the fan's optimum rpm, fitted inside a hollow high-speed shaft driving the compressor (unless an aft fan is used). A two-spool with its extra seals is a much more complex engine than a single-spool turbojet or centrifugal flow turboprop.
All of this was done and turbofans quickly became the norm, but it took a few years to develop and ramp up production.

I wonder if turboprop showed up any time earlier than turbofan. Neither the turbine nor the gear box of a turboprop is anywhere simpler than a small-medium BPR turbofan.

The shaft and the seals of a turboprop can be simpler - a gearbox (mature tech already) makes a single shaft design more reasonable. The fan isn't complicated, but it's got more blades that are thinner and have tighter tolerances than a propeller.

@user3528438, a gearbox isn't simple, but even back when the first turboprop engines came out, it was a mature, well-understood technology.  In contrast, a two-spool jet engine used highly experimental technologies.

@user3528438 the Vickers Viscount, the first turboprop airliner, first flew in July 1948, just over a year before the first flight of the dH Comet

@Therac But to make a small BPR turbofan you need neither a gearbox nor two spools, (nor a large diameter fan). I guess RR chose not to do it that way because performance wise that will be too similar to a turbojet and simply doesn't worth the resources.

@Therac conceptually, the fan isn't mechanically complex, but designing an efficient fan which doesn't stall, doesn't have aerodynamic flutter and/or vibration problems, and can withstand bird strikes, is far from simple - plus the requirement for a containment system if it does fail. The early fan designs using relatively simple solid blades, and snubbers between the blades to control vibration, wouldn't scale up to modern large turbofan engines. A propeller was much simpler to design than a fan, and experience with props dated right back to the Wright brothers.

@alephzero: In general I agree with your comment, but prop experience is much older: Lots of powered airships had them way before the Wrights.

No need for a two-spool engine if you're willing to go with a pusher turbofan.

Did they not fully realise the potential benefits at the time/thought they wouldn't work well?

Turbofans and turboprops were already studied in Britain and Germany during WW II, but at that time were technically not yet possible. Their benefit was indeed clear, even though the lower noise of turbofans was less of a consideration than it is now.

or was it just very difficult to build a jet with useful excess turbine power (to drive the fan) at the time?

Yes, that was the reason for the first jets. The pressure ratios of early jet engines did not allow to add a fan, and also the added complexity of a second spool was seen as not economical.
However, the next generation of jets could have supported a fan but didn't because they had been developed for fighters. Here the goal was to fly as fast as possible, and that needs low frontal areas and high exit speeds. Both would have suffered if a fan had been added. Civilian use of jet engines was initially restricted to the Comet, which buried its engines in the wings and could not accommodate a fan. Same for the V-bombers. Only Boeing's bombers could have benefitted from turbofans, but aerial refuelling made their low engine efficiency less obvious. Note that the Pratt&Whitney J-57, which was one of the first two-spool engines, was a pure turbojet for the first eight years of its life.
Only when the market volume of jet transport aircraft grew to a sufficient size in the late Fifties did the engine companies start to add fan stages to their existing engines. The first was the RR Conway, which had been developed for the V-bombers and had a bypass ratio of only 0.25 due to the diameter restrictions of the existing designs. The J-57 turbojet was developed further into the JT-3D turbofan in 1958 when the range of the first, turbojet-powered, Boeing 707s and DC-8s turned out to be marginal for transatlantic flights.

Yet early jetliners first used turbojets then only very low-bypass turbofans, despite the existence of turboprops

There were civilian turboprops for short range flights, but the speed advantage of jets meant that all new types developed for more than regional flying had to be jets in order to be competitive. Note that the latecomer in the trio of Boeing 707, Douglas DC-8 and Convair 880 used a smaller fuselage diameter and four Starfighter engines in order to gain a speed advantage over the other two.

Because, when the first jetliners appeared, there weren't any turbofans available to power them.  The first jetliner, the de Havilland Comet 1, first flew in July 1949; the first turbofan, the Rolls-Royce Conway RCo.5, wasn't available in a production-ready version until August 1955 (although, as it turned out, the first version of the Conway to power a jetliner was actually the RCo.10, which was itself rapidly superseded by the RCo.12).


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What preflight passenger briefing items are required under Part 91?
I have been taught to use the SAFETY acronym, mostly for checkride purposes. Which of those items are explicitly required (give references please) to include on a PAX brief for non-compensatory pt 91 flight?
My most recent DPE told me that only the seatbelts and the emergency exits are actually required. I found in 91.107(a)(1) the seatbelt requirement but cannot find the emergency exit req. He is right? Are there others?
Note: this scenario is assumed excempt from 91.519.

I’d say he is right. There is nothing in the AIM or any ACs that I am aware of.  But I always make sure that passengers know how to open the doors. Not a big deal on Cessnas since they work like car doors, but on Pipers with the overhead lock, it could be a problem in an emergency, especially since they only have one door and it is not next to the pilot.

There are regulations, and then there are best practices.  My personal checklist covers, seat belts, shoulder harnesses, egress (doors, baggage doors, etc.), quiet during TO and LNG, pointing out traffic conflicts or safety concerns, inadvertent window openings, motion sickness procedures, etc.  Should one ever have something go wrong, there will be lots of lawyers and FAA investigators looking to blame the pilot.

As far as I can find, the only briefing requirements in part 91 are for large or turbine powered aircraft at 91.519 (which you assume exempted from), the one you found along with the seatbelt requirement at 91.107(a)(1), and 91.1035 which deals with fractional ownership aircraft.
An AOPA article on briefings also states that:

Operation of seat belts is the only FAA-required briefing item.

The only references to emergency exits under part 91 is 91.607 which pertains to passengers for hire. It only mentions the requirements for the number of exits and does not mention passenger briefing.


text:
How to calculate density altitude using this chart?
I am trying to use the chart to calculate the density altitude. The question is : (Refer to figure 8 in your Test Guide book) If the current temperature is 90°F, PA 2000’ and altimeter is 29.70, what is the density altitude?
I did 2000' plus 205 since that is the pressure altitude convertor factor. 29.70 is 205 on the chart.
So I used 2205 and went on the chart and used 90 Fahrenheit and then got an answer of 4,600 feet. I feel like I am doing something wrong like there is an additional calculation before doing it.

If you are on a pressure altitude of 2000 ft, that is already assuming a standard altimeter setting of 29.92".
First you find the intersection of the line representing 2000 ft pressure altitude and 90°F. From there you read the approximate density altitude on the vertical axis. This will get you to ~4400 ft.

If your barometric altitude were 2000 ft at an altimeter setting of 29.70", then you would first need to calculate your pressure altitude. This would be done using the table on the right. The correction value of 205 needs to be added to the barometric altitude to convert to a standard setting of 29.92". This would yield a pressure altitude of 2205 ft.
Now using the density altitude chart again, you would find a density altitude of approximately 4600 ft.

Thank you so much! I guess I was heading in the right direction.


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What is the aviation term for what the navy calls "fire control system"?
The weapons systems, including computers etc, in a naval ship can be referred to as the Fire Control System. I'm not sure if FCS would be considered an appropriate acronym for fire control; I'm using to hearing FCS as Flight Control System.
Is there a specific aviation term equivalent to Fire Control System? If so, what acronym does it use (if any)?

The answer to this would depend on which armed aircraft one is flying.  Are you specifically asking about the system the is integrated with a gun, with missile systems, or both?

It is still called the (Airborne) Fire Control System; found about it here from a Google search.
Well, the acronym for (Airborne) Fire Control System is FCS (As well as Flight Control System) according to a military acronym website

@sanchises Fixed it.

When searching for 'aircraft fire control system' on Google I was mostly getting results for Anti-Aircraft Fire Control Systems on naval vessels, but being more thorough now I see that there are some sources that refer to the weapon systems on aircraft as a Fire Control System as well. That does still leave the other part of the question unanswered - is there an acronym for Fire Control System? The obvious one would be FCS but that would conflict with Flight Control System


text:
How is the serviceability of VOR unit part CI 215 checked?
In the VOR receiver we use two antenna, which are joined with the help of an assembly (part # CI 215), then the output of these two antenna is fed to RxR.
My question is: how do we assure the serviceability of this unit (CI 215)?

Means coxial cable inner lead should short with outer OR Not.
Both antenna Fitting point short with each other OR Not.
coxial cable inner lead should short with both antenna Base fitting point OR Not.

I assume you are referring to a VOR receiver of a specific model, you might want to add the name of the model / part number.

If the VOR receiver receives a signal, then the antenna assembly is working.

Problems with my aircraft is, On VOR ferquency To/From flag dit not come remain NAV flag.

Not sure I understand the problem, and the question. Can you confirm you suspect the receiver is not fed with the antenna signal (possible problem with the antenna array, or the coaxial cables), and you are asking only about the wiring between antennas and receiver?


text:
How does the energy and weight efficiency compare between a battery/electrical propulsion and combustion engine?
For the same total energy and power output would a battery powered vehicle be lighter than one which uses a fuel burning engine (Currently)?
This question is intended for helicopters.

Downvote because the answer is so obvious.

@PeterA.Schneider Not fair.  Answers are only obvious when you know them.  How many pilots and engineers, let alone folks not in aviation, understand energy densities?

@Simon Well, there are many questions (and their answers) on the Aviation stack which I truly enjoy. There are aviation experts here which provide insights and information not readily available to the general public. But the energy density of batteries vs. fossil fuels is not one of them. First of all the answer is trivially googleable. Second, the answer is visible in everyday life, starting with the limits of electric cars (which don't even attempt to fly!). Did the author never change a car battery which is as heavy as all the gasoline in the tank but sometimes can't even start the car?

@PeterA.Schneider Well, if Googlability was a criteria, then I guess somewhere between a third and a half of all questions on this site are not welcome.

@PeterA.Schneider 's statement is not so obvious to me, an electric powered vehicle can be lighter than a vehicle using fuel with same power. Indeed fuel (45 MJ/kg) is more efficient than Li-ion battery (0.5 MJ/kg). But this is for fuel only. Compare also the weight of the engine, and the rest of the infrastructure. There is a limit, for small power, where this statement is not true anymore. If a minimum power is specified, the statement may be true, but the question doesn't specify (now) a limit. So +1 to this question.

Poorly. That is a reason why we don't see battery powered aircraft beyond some experiments. The specific power is still much worse for batteries.
Energy storage:

Jet fuel has specific energy around 43 MJ/kg and energy density around 35 GJ/m³[1].
The best mass-produced rechargeable batteries to this day, Lithium-ion polymer, the best kind has specific energy 0.95 MJ/kg and energy density at most 2.6 GJ/m³[2].

That is more than an order of magnitude difference in favour of hydrocarbon fuel (almost 2 orders by weight, one by volume). Even taking into account the 2-3 times higher efficiency of electric engines (which do not suffer the theoretical efficiency limits associated with heat engines) hydrocarbon fuel is an order of magnitude better by weight.
Engines:

According to [3], the power-to-weight ratios for electric engines and gas turbines are comparable (in the range of 5-10 kW/kg).

Electric engine won't give any advantage to electricity.
Well, the difference in energy density is so huge that even large difference in engine weight wouldn't help much anyway. For models, electricity is OK as due to the square-cube law they have a lot of power and lift and small-scale heat engines are difficult. But for aircraft carrying humans, batteries are very far from usable.

@Federico, 43/0.95 = 45, which is more than $10^{1.5}$.

oh, sorry, was looking at volume. thanks for having updated your answer.

And it gets worse with flight time. While fuel is consumed and the aircraft gets lighter, batteries will have the same mass, regardless how much energy is left.

Unless solacells can provide some of the power.....

Everything is limited by the second law of thermodynamics.  That's why it's a law.  Perhaps you're thinking of Carnot efficiency?

@J..., that's the same thing. Second law of thermodynamics says what you can't do with heat and since electric engines don't work with heat, it has no consequences for them.

@JanHudec No, the second law says that entropy in a closed system increases - this applies to electric motors as well as to combustion engines.  Anything that isn't limited by the second law is either a perpetual motion machine or a free energy source.  The efficiency limit of a combustion engine is best described by the Carnot cycle.

@IanRingrose If you put sixteen 777-200s in a square of 4x4 (wingtip to wingtip, nose-to-tail) and drew an unbroken square around all sixteen aircraft, you would have a solar cell that, if it was 20% efficient (typical for a good cell), would produce about half the power of one Trent-800 engine (in full sunshine, pointed at the sun).

I was involved in a project to electrify small aircraft. If the best batteries available would increase their energy density by 3 to 5 times of the current capacity, then we are talking about equal range to an AVGAS powered small aircraft. This takes into account the much higher efficiency of the electric motor


text:
Can I fly a glider in Europe on a Canadian Glider License?
I currently live in Canada as a Permanent Resident and I recently received my Canadian Glider Pilot License. I will be going to Europe soon so I was wondering what I have to do to be able to fly a glider there. I will be going to the Netherlands if this makes any difference and I still have my Dutch Citizenship. Do I have to convert my license or just do some extra training?
Also would it help if I got a FAA glider certificate/endorsement on top of my TCCA license?

Hi Laurens, welcome to aviation.sx!

Make sure you have the right medical too. US pilots who fly without a medical must get a medical to fly in France.

I currently have a Cat 1 Medical in Canada.

If all else fails try finding a local gliding club and see if you can fly with one of their members.

In general, it is difficult to convert to EASA. You can find more information here however I suspect the easiest method will simply be for you to phone them up.
I think the bureaucracy will probably be more hassle than it's worth. It's not like Transport Canada where you can just get a FLVC (Foreign Licence Validation Certificate) or whatever (that was a pleasent suprise after our CAA in the UK!).

TC is short for Transport Canada, and FLVC (Foreign Licence Validation Certificate) is something they issue. I would say both are fairly common, especially for Canadians.

Sure, I'll edit :)

Before EASA you could fly a German-registered glider on almost any nation's glider credential(s). I did it for a year before getting a German SPL just because. No idea if this still exists.
You can certainly fly gliders w/ an instructor. If you are a visitor you're highly unlikely to get to solo a club's glider. Not sure about NL, but in Germany there are places that function as commercial training facilities though technically clubs w/ daily membership.
If the instructor is willing to let you solo, you could simply solo as if you were a regular EASA Land student glider pilot.
I got a British (BGA) Gliding License super easily by email and credit card. I flew several times at really nice places in England and Scotland.

If the glider has Canadian registration, you can fly it anywhere in the world. If it doesn't, it depends on which registration it has and in what country you have to fly. Each country is different, and the further west you go in Europe, the more anal[-retentive] the bureaucrats. UK is the worst, as you go east in Europe it gets easier and more laid back. I have an N-registered glider in Germany so I can fly it any time I want, but I can't fly a D- glider. Unfortunately, to get the correct answer for any country you have to navigate their regulatory bureaucracy, which can be pretty thick in many places.

Welcome to Aviation.SE. Please check the language used in your answer, especially the part about bureaucrats is inappropriate for this site. Can you also provide sources and citation for some of your remarks? Otherwise it's rather just an opinionated comment, but does not really answer the question.

@ymb1 I know what he meant, doesnt make it appropriate language.

As @SentryRaven stated in his second answer, there is nothing vulgar or inappropriate about the term. Move on.
The answer is the last sentence. It means "You have to do your own homework in each country."


text:
Where can I find a list of internal airline designations for various aircraft/cabin configurations?
Is there any way to tell what the internal designations of an airline mean? Looking at Delta's seat maps for the 757, Delta has a couple of kinds, but what's the difference?

Boeing 75A
Boeing 75M
Boeing 75S
Boeing 75G

I know they're slightly different configurations for different routes, over water etc, but can what type is used on what route/configuration or what the designations mean be found somewhere?
Is there a place where these are defined for airlines of the world?

Internal designations are usually used to differentiate various seating configurations or differences internally for an airline. From your example, all those codes are various codes for the 757-200 (at Delta, they code the 757-300 as a 75Y). You can often look up various seat configurations at Seatguru.com, a site that generally tracks seating information on airlines. Additionally, all would be sent to ATC using the ICAO type B752.
The reason for these variations can be broken down to a couple areas of differences. One is exit configuration; in the same type (757-200), there can be 2 or more different configurations of emergency exits (combinations of doors and windows, and various types). If you have a fleet that mixes these, you'll often have to have different seating configurations for them. Another differentiation could be whether there is a subfleet equipped and used for ETOPS flights. There can be other reasons an airline could want a subfleet and a different cabin. Also, if you're in the process of upgrading a fleet, you might want to create codes to differentiate them so you know which is going where and be able to plan your revenue models around that.


text:
How to identify a specific 707 variant?
I recently saw a very interesting documentary film made in 1980 about a narrowly-averted airline disaster in which the flight crew were incapacitated and a retired pilot, with the help of other crew members and experienced ground controllers, managed to land the plane safely.
The aircraft in the film is a 707, though I was puzzled by some of the interior shots, in which I thought it looked much more like a DC-8 (it's been a very long time since I flew in either). The footage of the cockpit was actually quite baffling, and I really didn't recognise it at all.
What exterior and interior features would help identify the particular 707 variant?
The documentary is apparently quite famous, so perhaps someone even knows the tail number. It's simply called Airplane! - very informative, and I learned a great deal about the aviation - highly recommended.

The DC-8 and B707 are different aircraft families.

Did you also notice that the 707 sounded throughout the entire, uh, documentary like it had propellers?

I saw that documentary once while in a turkish prison.

I'm not sure if you want to identify the aircraft used in that particular program or in a general sense; if the latter is the case, do try airlinercafe.com

Don't know about the interior shots but for the exterior:

In-flight shots use a model, some ground shots use N6723, c/n 18989/492, a TWA 707-131B.

This is from the awesome website IMPDb Internet Movie Plane database

Apparently, my browser does not like that website, or the other way around.

@SMSvonderTann You can try their main page at imdpb.org. it's a very cool site and folks on here could probably help them fill in gaps


text:
How is the cabin pressure maintained inside aircraft?
When aircraft is at high altitude, atmospheric pressure will be too low but inside aircraft pressure is maintained such that it is comfortable for crew and passengers. But how is it maintained?
As I understand that at higher altitude pressure difference can be too large which can deform the plane than how this is handled.

Once the airplane is closed up, all doors and windows secure, pressurization comes down to the following:

how much air is being pumped into the cabin?
how much air is being let out of the cabin?

Air enters the cabin (on a jet powered airplane) through bleed air taps in the high pressure compressor of the engines.  This air is routed through air cycle machines, or "packs" as they are commonly referred to.  The packs condition the air for temperature and humidity (typically taking all of the water out of the air) and then pump this air into the cabin at high pressure such that it is at the proper temperature after expansion to cabin altitude.
Air exits the cabin through outflow valves, which are typically at the rear of the pressure vessel.  These can be manually or computer controlled and will open and close to regulate how much air leaves the cabin.
To provide constantly fresh air, the incoming air from the packs is constantly supplied and as such the cabin pressure is almost entirely regulated by the outflow valves.  These valves have to be able to exhaust more CFM of flow than the packs can supply in order to let the cabin climb and when providing less CFM
than the packs the cabin will descend.
The cabin altitude will be managed so that:

it stays below a maximum pressure differential with the ambient atmosphere outside,
cabin altitude rate of change is limited,
maximum cabin altitude is kept below 10k feet (and typically below 8k feet) such that supplemental oxygen is not needed.


text:
What DAL are Humans?
I am no expert at the Development Assurance Process, but hypothetically speaking what Design Assurance Level (DAL) would a human pilot receive if they were classified as one (i.e. DAL A, B, C, D, E). It seems absurd to try and answer this question since the plane can't even leave the gate today without a pilot or pilots, but I am hoping to learn more about the DAL process thru any answers received.

"It seems absurd to try and answer this question..."  I agree.

Come on, at least tell us what DAL means.

@quietflyer see the DO-178 specification for definition of DAL (or quick summary here https://en.wikipedia.org/wiki/DO-178C)

I'm oversimplifying a bit, but more or less the DAL level of a piece of software is decided by asking the question, "What is the worst that could happen if this software fails to operate as designed?"  Note that "failure" in this case could mean that it simply stops working, but it could also mean that it does the wrong thing, like pitch up when it should pitch down. So for example,

The software that runs the in flight entertainment system.  Worst thing that could happen has no safety effect whatsoever, therefore level E.
The software that runs certain cockpit indicators, like the engine vibration display or engine oil temperature display: Usually considered a "major" safety effect, but not more than that, so level C.
The software that controls the engine thrust: if that fails to set the thrust according to the pilot demand, that could result in the plane crashing, so "catastrophic effect", thus level A.

It may be a little subjective, but I'd say that if we consider the pilot to be "software", if the pilot fails, the worst thing that could happen would be "Catastrophic" condition, and thus I would label the pilot as level A software.
Edits: clarified that DO178 applies specifically to software

"In flight entertainment system. Worst thing that could happen" - Swissair 111?

@Zeus Swissair 111 was a wiring problem. In that particular case, that wiring happened to be for the IFE system, but the exact same problem could in principle have developed with any other wired system. Therefore, if anything, it stands to reason that you should apply the same reasoning at least to any wiring carrying significant voltage or power.

@Zeus interesting, I was unaware of that one.  In any case, DO-178C only applies to software.  There are separate regulations governing hardware. i.e. flammability and electric wiring don't fall under DO-178C.  I'll edit the answer to be more clear on that.

@aCVn, true, of course, but it shows that it's dangerous to think (and state unconditionally) in that vein. The system has to be considered as a whole. We can even speculate what could happen from a software bug, assuming correct wiring: say, an infinite loop causing excessive power consumption for all seats at once -> failure of circuit breaker -> shedding of more critical load.


text:
What was the first control room--used for flight test--established in aerospace history?
The control room for space missions in Houston is probably the best, most widely known example of a control room.  It is an operations center used in support and execution of flight and space missions.
Historically, what was the first instance wherein a test team used a control room or operations centers for flight operations support?
This question answers why a control room exists, and this question answers who gets to be in there.

It depends on how you define support; voice only or first use of radio-transmitted telemetry?

How about providing an answer or examples of both?

This article would seem to suggest that use of radio telemetry started being used sometime during the mid-fifties.

Then you reduce these rolls of oscillograph readouts and learn to read the data, you will know more about what happened during a flight test than the pilot, the engineer and the designer. These rolls of paper are like novels. It is up to you to get the meaning, then sense the plot and determine whether flight objectives were satisfied.

The quote above is from chapter 5 of "Failure is not an option" by Gene Kranz, describing his first introduction to flight testing in 1954 when he started working for McDonnell Aircraft. As the quote would indicate, data processing happened after the flight was over and that there was no kind of control room monitoring the progress of the flight. An exception was for incidents such as one also described by Kranz, where a B-52 had lowered a missile from the bomb bay in order to fire it. The missile would neither fire or retract, but as the bomber had enough fuel the engineers had enough time to discuss various options.
I suspect that the first form for flight operations support would have been the chaseplane, although this is outside what you are asking.
The next closest thing I've heard about is Mercury Control, which was used to monitor and support flights during the Mercury Program.

I think you're looking way too late here - the concept of a control room was around well before that. As a top of my head example - RAF Fighter Command as an example of real time tactical operations using radio telephony, RADAR and other sources to command and control the RAF's fighters.

But were RAF fighter command use those to support test flights? The impression Kranz and others make is that real-time monitoring was a brand new concept when they started with Mercury, but that they built that based on lessons learned during aircraft testing.

What about V-2 development and testing? After all, those were the same engineers who went on to work on the US space program. The flight duration was shorter, however.

V-2 might be the first, or at least one of the first. According to a reference in Wikipedia some telemetry was received from a remote site, but it doesn't say what. This also makes sense, as a big difference between testing aircraft and rockets is that the aircraft usually lands intact in one piece.
If you know you're getting the hardware back, why bother adding lots of extra weight to transmit data when it can simply be recorded (in much more detail). Rockets on the other hand usually try to disassemble themselves into as many and as small pieces as possible, so recovery is less likely...

That depends on how far you are willing to stretch the definition but a case can be made for as far back as 1903/1904. The Wright Brothers used a small building to support the initial flights of their early powered aircraft. It can be seen here in this photo although this also makes a good candidate for the first hanger in history.

A case could be made that Wilbur Wright also served as the first ground support individual in the history of powered flight.

(December 17, 1903) ...Wilbur Wright running alongside to balance the machine, has just
  released his hold on the forward upright of the right wing...
  starting rail

The ability for an aircraft to communicate with the ground and have remote operation coordination occur dates back to 30's when radios became popular on aircraft. So its conceivable that ground support for testing was a reality around then.

In his 1935 speed record flight Howard Hughes had a ground support crew, timing, filming and coordinating the flight. While they were not in a room per say you could call them a kind of mission control. This is the best image I can find of the group:

As far as standardizing it, developing the best methods, and honing the operation of, NASA gets most of the credit for modern mission control.

I think it's a "stretch" to use the First Flight as an example.  Wilbur did little more than watch.
The Howard Hughes reference is good, but was it the first?
And if you claim NASA gets the most credit for standardizing, I would like a reference.  But as the other answers in this thread suggest, NASA employees brought their experience with them from previous employers.


text:
Who was the first flight test engineer?
The discipline of Flight Test Engineering has existed since at least 1903, but when did engineers identify their work, distinguish their work from other disciplines, under terms that would mark it as Flight Test Engineering?
This may take the form of someone who first used the term flight test engineer or flight test engineering.
Google demonstrates a history of the terms:

Image credit
NATO describes the history of its flight test organization this way:

Soon after its founding in 1952, the Advisory Group for Aerospace
  Research and Development (AGARD) recognized the need for a
  comprehensive publication on flight test techniques and the associated
  instrumentation. Under the direction of the Flight Test Panel (now the
  Flight Vehicle Integration Panel) a Flight Test Manual was published
  in the years 1954 to 1956. This original Manual was prepared as four
  volumes: 1. Performance, 2. Stability and Control, 3. Instrumentation
  Catalog, and 4. Instrumentation Systems. As a result of the advances
  in the field of flight test instrumentation, the Flight Test
  Instrumentation Group was formed in 1968 to update Volumes 3 and 4 of
  the Flight Test Manual by publication of the Flight Test
  Instrumentation Series, AGARDograph 160. In its published volumes
  AGARDograph 160 has covered recent developments in flight test
  instrumentation.

What is the difference between a test pilot and a "flight test engineer" especially in the context of early aviation?

Are you asking for countries where English is spoken?

@mins No. But I don't the translation in any other language.

@RonBeyer probably very little, but self-identification as an engineering professional is part of the nuance of the question.

I'm sure the Wright brothers would have identified themselves as such had the term been in use at that time. The actions they took (building a wind tunnel, multiple wing designs, glider trials, etc) would absolutely have qualified them as such. Even before that back during balloon flight by the Montgolfier brothers in 1782/1783 could qualify as basic test flight engineering.

@RonBeyer I agree with your assessment, but my desire is to find out "when" the term first came into the lexicon of the professionals in this field.

In small  industries people can wear many hats and may not call themselves a particular specialty.  Self identification of flight test engineer is partly from the expanding depth of specialization, but also expanding industry needs  So largely you would be looking for the point where there are enough large players in the industry to invent the new title. A similar situation is prospector vs geological engineer , a specialization that played out in the old west without a hard fallover date, basically it relates to when universities and schools  start teaching it as a specialty.

@crasic A reference that shows the first use of the term would be a step in the right direction: "This may take the form of someone who first used the term flight test engineer or flight test engineering."

@RonBeyer in the modern sense at least, an FTE is a non-pilot position.  They're the ones running the computer consoles in the back.

@JohnK Maybe for large aircraft like 60+ seat airliners, but smaller aircraft don't have dedicated test engineers on board during a flight, so the pilots have to act as a flight test engineer. I'd say the majority of FTE's have at least a PPL, at least it has been a requirement of all the FTE job applications I've looked at.

The Michigan Alumnus written on Aug 13, 1938 refers to Clarence Johnson's experiences as a  Flight Test Engineer. He graduated with a MS in 1933 so presumably became an engineer at that time.

How on earth did you dig that up?

@JohnK Doesn't everyone read eighty year old alumni newsletters?)

Presumably is not an answer. And I have references with earlier answers.

@MarkJonesJr. You can answer your own question. Put a better one up since you didn't like the only answer you got in two years. 1938 is the answer if you object to 'presumably' 1933.


text:
Why do private jets such as Gulfstream fly higher than other civilian jets?
I heard in a TV show that private jets such as Gulfstream can fly at about 50,000 ft, higher than other civil jets.
Is it an aerodynamic reason (lighter aircraft to be sustained in a less dense air) or does it depend on the fact that a smaller fuselage is more resistant to lower outside pressure at a higher altitude?
I think that – if possible – flying higher, in order to fly faster and with lower fuel flow, is always better. Am I correct?

Jet engine efficiency does not increase above the tropopause, at around 11km, due to the way temperature behaves beyond it, see here: https://aviation.stackexchange.com/questions/12541/why-are-many-jet-aircraft-designed-to-cruise-around-fl350-370

Re, "higher is always better:" If that were a simple truth, then long-range flights all would be in space planes.

@SolomonSlow Technically, the longest long range flights are: Any flight around the earth one or more times are in some kind of space plane.

@VolkerSiegel It seems odd to call a plane with an service ceiling of 51,000 feet and rather limited maneuverability in a vacuum a space plane.

There are several reasons and none of them really have to do with a private jet being more aerodynamic than a commercial jet.
Private jets have better power to weight ratio than commercial jets so it is easier to ascend to greater heights. Commercial jets also carry more fuel which needs to be burned off to lighten the plane to go higher.
It’s easier to pressurize smaller cabins than larger ones (this is a significant factor for commercial jets).
Flying higher is usually better for avoiding weather that can even affect commercial airliners in the 30ks.
There is much less traffic up above 40k and it allows the private jets to fly more direct routes at a faster speed and more safely as well.

You forgot the bragging rights :-)

They are designed for higher cruise speed, which explains the higher power/weight ratio. Add to that a lower wing loading for operating out of regional airports. The higher cruise altitude is just a consequence of the desire to fly faster and from more airports.

There are three main factors that let corporate jets with 50000+ ceilings get up that high.
First and most importantly, they have very large wings.  Partly this is because of the need for fuel volume for range requirements, but it's mainly that you need the extra wing area to be able to fly that high with a reasonably efficient indicated airspeed at 50000 ft (500 kt TAS is only 250 kt indicated at FL500) at reasonable weights.
Next they use turbofan engines like the BR700 with lower bypass ratios than airliners because a lower bypass ratio is somewhat more efficient in very thin air.
Lastly, they operate with higher cabin pressure differentials, 9-10 psi rather than around 8 psi to get an 8000 ft cabin at 50+ thousand feet.
The structure has be a bit heavier to handle the pressure differential, but not that much because they are designed for way less cycles than an airliner, 3-500 cycles per year vs 2-3000 cycles for a regional airliner.  If you took a corporate jet as-is and put in a regular interior and started running it 4 or 5 flights per day flogging it like a mule, it'd start breaking down in short order.  Conversely, corporate aircraft created as conversions of regional airliners designed for 3000 cycles per year will run pretty much forever in the light duty corporate role.

The structure does not necessarily need to be heavier. Since the diameter of the cabin is smaller, the square-cube-law helps you out, i.e. you can have higher pressure differential using the same materials and wall thickness.

Not necessarily; Global's tube is the same diameter as the RJ's.

Agreed, but that's a large long-range business jet compared to a small regional airliner.

I think the other important factor that the others haven't addressed is Mach number.
The speed of sound is doesn't change much with altitude (it actually drops slightly due to reducing temperature) and if a plane gets too close to it, shockwaves start to form in areas where the airflow is accelerated (such as on the top of the wing) causing a lot of excess drag.
By flying higher, a plane gets a higher ground speed and true air speed for the same indicated air speed (IAS). However the limit is when the true air speed (TAS) results in a Mach number that starts to cause shock waves. To go higher, a plane has to reduce its indicated air speed to maintain the same Mach number. Therefore the limiting height is when IAS approaches stall speed at the same time as the Mach number limit is reached. (see Coffin Corner)
Now private jets are usually designed to use smaller runways than big airliners, so they have a slower stall speed (due to relatively large wings for their weight). Their maximum Mach number is similar (an easy way to tell is that the angle of wing sweep is similar) so the height at which minimum IAS = maximum TAS is higher.
EDIT: True air speed (TAS) is the actual speed of the air, similar to the aircraft's ground speed (after allowing for wind). Indicated air speed (IAS) is the effective aerodynamic air speed, or "how much force you'd feel if you stuck your hand out the window". IAS falls relative to TAS as the air gets thinner.

How much faster would a Gulfstream at 50,000ft be able to fly a route compared to a commercial flight at 35,000? (say SFO-JFK?)

@Aaron, no faster. It's flying at the same Mach number and thus approximately the same true air speed and ground speed. The thinner air means that the indicated air speed is lower, so there's less drag and the engines don't need as much thrust.

I am quite surprised that no-one mentioned safety as a reason in the previous answer. In 1996, the FAA adopted Amendment 25–87, which imposed additional high-altitude cabin-pressure specifications, for new designs of aircraft types.

The standards adopted as Amendment 25-87 pertain to operation of subsonic airplanes to a maximum altitude of 51,000 feet, although many of the requirements addressed therein relate to operations at lower altitudes (below 41,000 feet) as well.

Typical systems to be considered are engine bleed air systems, air conditioning systems, power sources, outflow valves and their associated control systems, etc. Failures or a combination of failures which expose the occupants to: (1) cabin altitudes in excess of either 25,000 feet for more than 2 minutes, or (2) cabin altitudes that exceed 40,000 feet for any duration, shall be shown to be extremely improbable (see Glossary).

Airliners altitude are thus limited by safety regulations in case of rapid pressure loss. On modern airliner a major engine failure might lead to a quick depressurization and even passenger sucked out of the window. In this case you have to lower the altitude below 25'000ft in less than 2 minutes which means a decent rate of approximately 10'000ft/min. To avoid any structural overstress, you can't have much higher decent rate than that thus limiting the maximal altitude you can fly. The 40'000ft absolute limit is not unbreakable as the A380 was certified to fly at 43'000ft.
On the other end private jet have the engines mounted on the aft fuselage, and behind the pressurization bulkhead. Which means that fast depressurization in case of an engine failure is totally unlikely and they can therefore bypass this rule. Furthermore they operate at higher cabin pressure which means in case of an air pressurization system failure, they will have a bit more time to descend to lower altitudes. Therefore they can use higher altitude up to 51'000ft to avoid traffic and increase fuel efficiency.

wait so why does engine location affect depressurisation? I thought it was simply a matter of using bleed air for the cabin

In case of engine uncontained engine failure some part might puncture the airframe. The private jet pressurized cabin ends before the engine location thus mitigating the risk of depressurization.


text:
Can a short in instrumentation wiring create enough 'spark' to ignite unvented fuel fumes?
Although this question relates directly to the NTSB findings on TWA Flt800, I was always puzzled and dissatisfied with the reason that failure in the instrument wiring harness inside the center wing fuel tank(CWT) caused the unvented fumes(a failure in itself) to explode; given that instrument circuits use microamps at 28vdc, resulting in less than .02Joules at best to transmit signals. A probe heater of course would be different, but that wasn't given as the reason. Are there any other instances where wires 'shorted' in an instrumentation circuit could cause sufficient energy to create a spark?
This question is not about "what really did happen" to TWA800; this is about whether or not the NTSB(National Transportation Safety Board) made a correct analysis in implicating Instrumentation Wiring.
Note

A joule (J) is a unit of measurement of electrical work or energy; 1 J
  is the amount of work done by 1 watt of power in 1 second. The power
  supplied to the FQIS components through the cockpit fuel gauge (0.02
  mJ) is about 10 percent of the minimum ignition energy (MIE)
  requirement (0.25 mJ) for hydrocarbon fuels, referenced by the
  American Petroleum Institute Recommended Practice 2003 (API 2003),
  Protection Against Ignitions Arising out of Static, Lightning, and
  Stray Currents, fifth edition, December 1991 Taken from here

there were other factors IIRC, including deposits on the fuel sensor that shortened the spark gap

Milliamps is not enough.  You need to know the voltage and many other factors such as the spark gap medium, the dielectric function of that medium, the spark gap size etc etc.  A very high voltage with tiny current (static electricity) will generate a spark as will a low voltage with a high current (try shorting a car battery).  28vdc at 250 milliamps will happily melt aluminium.  28vdc at a few tens of milliamps will happily create spark across a millimetre or so of a reasonably conducting medium.

I feel this is more appropriate for physics.SE (spark gap as function of tension, energy released as function of gap, tension and current, etc.)

I don't think there are only signal wiring in those harnesses.  There could be power lines to instrument sensors and other circuits in the bundle.

My initial research indicated that a minimum of 1 Joule(I(2)R) is needed to cause sufficient arc to ignite the combustible Center Wing Tank(CWT) gasses. Honeywell and Boeing stipulated that the fuel probe system at max carries .02J. CWT explosion definitely was the cause; instrumentation wiring(IMO) seems an inadequate answer.

@Federico To those who put this on hold....I'm not asking about the physics of what causes a spark, nor am I asking what temperature can Jet "A" ignite. I am asking whether an instrumentation wire (on an aircraft) can produce enough energy to ignite unvented gasses in a fuel tank-so those nerds who calculate quarks will likely send it back here....

@user2479 You have two questions there: can the wire spark, and can the fuel-air mixture ignite. Both are physics questions.

@user2479 After your edit, it has become opinion based.

@Farhan  I'm just using the information contained in the NTSB report(you can find it on pg.34). What's interesting after reading the whole report(ouch-the eyes), even they said it was a 'suggestion'-they had no further evidence. I have worked on A/C systems and have replaced malfunctioning wiring-yes, if you short out canon plugs or route wiring wrong-bad things will happen. I have yet to find where an instrumentation circuit caused a short that 'arc'd', but I could be wrong.

@Simon If you just use 28*.02, you end up with .58w, which isn't Joules, but an energy measurement just the same. The Board said >1J will cause an arc; the manufacturer said their system would produce <.02J. So, what's the answer?

the answer is that you have to compute the Joules, hence is a physics.SE question.

@Federico The question deals with aircraft, not physics. I am on the physics SE site also, and it's mostly made up of 'brainiacs' working out advanced calculus solutions to theoretical problems. They 'exclude' engineering answers, as they are about pure physics(ahem). This question is about "whether or not anyone has seen this...in an aircraft", NOT-how many Joules per coulomb does it take to ignite a petrie dish of Jet 'A'.

@Federico Close and delete this question if you insist, however, this question 'may' clear up some 'myths' about insignificantly tiny energy sources taking down large commercial aircraft. Notice, I didn't say what DID take it down, but unless someone here(not the geeks in the Physics Lab) can prove that small of an energy source in the adjacent area of the fuel cell could ignite the fumes, I'll stand on my hunch.

I feel it would be better if you would ppass by chat. Anyway, as it stands now, you are asking whether or not the NTSB made a correct analysis in implicating Instrumentation Wiring, i.e. to redo the analysis, an analysis that involves lots of physics' considerations and computations.

@Federico At the end of their lengthy report, they admitted it wasn't conclusive, but based on their probability model(1x10(-5)), they list it as the best scenario. What sticks out also is the Flt Capt.'s final comment on the CVR "...look at that crazy fuel flow on #4..see that", which also makes a compelling reason, although not totally proven one.

@Federico(finally) What was interesting was it seems like they were blaming the heat from the Air Cond. packs, which were adjacent to the CWT and vented hot exhaust next to the access door to the CWT, yet later they ruled it out.


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In an unpressurized cabin, why is the cockpit ambient pressure lower than the outside pressure?
As I study for my IRA written, this (VFR) question resurfaced and continues to irk me. The annotations cite the Venturi effect as the sole reason, but I don't follow.
From Wikipedia: "The Venturi effect is the reduction in fluid pressure that results when a fluid flows through a constricted section (or choke) of a pipe"
Do we consider the parcel of air in the cabin to be the fluid moving through the constricted section of some imaginary/abstract pipe? If so, why? Can anyone shed more intuition on this?

The air moves around the fuselage, which causes a localized increase in velocity around the aircraft.  This increase in velocity causes a slight decrease in the outside air pressure, which effectively sucks air out of the fuselage, making the cabin at a slightly higher pressure altitude.
The effect is localized.
While the reduction in pressure, and increase in pressure altitude inside the cabin is small, many GA aircraft display a 40 foot or so difference in pressure altitude between the alternate static port (inside the aircraft) and the pressure altitude with the factory static port.

Opening the air vents can change the pressure inside to be higher than outside. Barry Schiff gives a good demonstration in a Wonderful World of Flying clip. https://www.youtube.com/watch?v=eHnhRrbyEIE

The original poster may be helped by imagining the fuselage as a carburetor venturi turned inside out.  That is to say that the constriction is on the outside of the aircraft going through the air, rather than inside the cabin.

The opening and closing of air vents is a different effect, largely the introduction of ram air into the cabin.  A window could increase the pressure with a slip into the open window, and decrease the pressure with a slip towards the side with the closed window.

Diagram would help here.

Sorry, @Notts90, I am not a good artist.  If it helps visualize, think of a wing, somewhat symmetric along the chord, with "lift" on the top and on the bottom and in this example, on the sides.  That "lift" is created by a small pressure differential (Re: Bernoulli Principle) caused by a small increase in air velocity as the air moves around the fuselage.  I will see if one of this term's aerodynamics students would do a diagram with calculations for extra credit.

@mins I am sorry you are misled. The velocity of the air inside and the external flow are not really relevant. Where you are on the money is with your analogy to wings.  This problem is similar to where there is a port to the top of the wing, and when that port is open, air leaves the cabin to match the pressure at the top of the wing. In the question instance, the entire fuselage is an airfoil, and the differential in pressure is integrated over the fuselage through leakage to the cabin, and this results in a lower absolute pressure in the aircraft.

The molecules of air in the aircraft remain there, unless part of the effluent "sucked" from the cabin through leaks in the doors, etc. The acceleration of the cabin air due to the fuselage velocity changes is not a relevant factor in this problem. If the plane fuselage were 99% lead would the problem be different? The lead having substantially more inertia than air.  I am just trying to understand our dissonance.

The point is that inertia of the air inside the fuselage is not a factor in this problem.


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How dangerous is to take-off and fly with snow on the wings?
I have seen a video of a commercial aircraft taking-off with snow on the wings. I wonder how dangerous it is to take-off and fly with snow on the wings.

2004 article: Ice Contamination Accidents On Takeoff Reach Epidemic Proportions

On Skeptics.SE: Can a sandpaper-thick layer of ice reduce lift by 30 percent and increase drag up to 40 percent?

Pilots are allowed to commence a takeoff if the snow is not adhering to the wing and will blow off during the takeoff roll.
It is very difficult to just look at a snow covered wing and know for sure if the snow will blow off. A "hands on" tactile inspection should be conducted and even then you can never be sure about the level of contamination for the entire wing surface.   De-Icing is always the safest option.
To attempt a takeoff with snow and/or ice adhering to the wing is NOT SAFE.  You are gambling with your life, and the lives of your passengers and crew.
That being said, it is possible to do a reasonably safe takeoff with snow and/or ice adhering to the wing, but you will need a much longer takeoff roll, and slow rotation, and climb out.  If you were to have an engine failure you would probably not have sufficient performance to climb out over obstacles, or even remain airborne.
You are gambling that you have enough runway for your "seat of the pants" takeoff and that you will not have an engine failure, because you have no way of knowing the actual aircraft performance.
So again, De-Icing is the only safe option.

What if you're taking off from a runway that's vastly longer than the normal length necessary for a normal takeoff in the type of aircraft you're flying (say, a Twin Otter on a 10,000-foot runway)?

With a longer than normal runway a safe takeoff is possible if you use a higher than normal takeoff speed. BUT.... if you were to have an engine failure once airborne, you may not have enough performance to continue climbing or remain in level flight.

If it's just light snow, it should blow off. If,  however, there is ice under that snow, the situation becomes dangerous.
Air Florida Flight 90. It attempted a takeoff with snow and ice on the wings and engine inlets from Washington National, and went into the Potomac River, a combination of inadequate power from the engines, and inadequate lift from the ice encrusted wings. Out of 74 passengers, there were five survivors.
Air Ontario Flight 1363 crashed on takeoff due to ice on the wings, a failure to de-ice after the aircraft had been on the ground at a regional stopover. They couldn't shut the engines down, because the internal APU wasn't working, and couldn't de-ice with the engines running. Bad choice... should have stayed on the ground.
Three years after that, USAir 405 also crashed on takeoff due to ice on the wings, in the same type aircraft: a Fokker F28. In this case, a lengthy delay on the ground while waiting for takeoff let freezing rain build up on the wings.
After USAir 405, deicing procedures and regulations in the US were found to be inadequate, and overhauled.


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Why might the crew switch off air conditioning in the cabin of a passenger jet during descent?
I've noticed on some commercial flights (most recently today in an A319) that as the aircraft descends in preparation for landing, the air in the cabin gets hotter. I also noticed that this seems to be caused by the air conditioning being switched off entirely or to a lower flow rate.
Is there a reason for the cabin/flight crew to switch the air conditioning lower/off for landing?

See also https://aviation.stackexchange.com/questions/22285/is-there-an-faa-regulation-to-switch-off-the-air-conditioning-during-take-off/22286

@MJeffryes Also possible duplicate: https://aviation.stackexchange.com/q/8408/65

@Farhan That question is about takeoff. I'd expect the answers could be similar, but it is definitely a different question.

It could be they are configuring for max thrust in case of a go around. It might also be the a/c packs are running less efficiently due to engines being at flight idle during descent.

The cabin must be re-pressurized when descending, pressurization is done with the packs, so packs should be ON, at least one. In addition, temperature would rise if packs were OFF (due to passengers heat and outside temperature increasing too).

TomMcW got it. The quick answer is that the engines are at idle = less bleed air = packs have less capacity.

Systems vary.
Transport aircraft and larger biz aircraft have airpacks handling the AC and at flight idle power settings there is less bleed air, and pressurization is a priority.  The packs simply do not have full power when in a flight idle descent on many aircraft.
The aircraft I routinely fly have separate radial compressor driven refrigerant systems, and have large heat loads from onboard electronics and laser systems.  They have to remain cool, and normally both AC systems are left on until just prior to engine shutoff after taxi.  If it is a hot environment, we use the APU and keep the AC on after engine shutdown.
Once again, the situation is different in different aircraft, and is dependent upon the cooling system, the power plants, the installations and the design and operational goals.
Addendum #1: Some aircraft APUs can provide bleed air for cabin pressurization, but this is not universal.  Also, some aircraft, such as the 787 do not use engine bleed air for pressurization, and rather use electrical compressors.  The flow of pressurized air is rather constant in flight.

Do you know if the latest incarnations of commercial aircraft (A380, B787, etc.) have a way of driving the AC and maintaining cabin pressure without engine thrust?

B787 is bleed-less, so no air from the engine is used for AC. All others, especially the B7478 (nearly same engines, GEnx-series from the B787) and A350 (the GEnx is an option there, but Airbus probably decided not to invest so much effort to redesign everything). Of course during flight the default energy-source are the engines, as long as the APU is not used.

My answer is purely speculation, and someone with knowledge should chime in, but my guess is that the APU can power AC, but not pressurization.  However, I do know that the 787 uses electrically powered compressors, so I would therefore expect (another assumption) that pressurization is possible with the APU running. So after the fans quit, the descent may be with a pressurized cabin (and electricity to power the control surfaces).

@Peter we were typing at the same time, but it appears that the 787 relies on electrical power for compressors to pressurize and also to run refrigerant based environmental control.  I wonder how cabin heat is managed...through electrical resistive heat?

There's a misunderstanding here. A jet pressurized aircraft does not have a separate pressurization and air conditioning system. The cabin pressure controller only controls the outflow valve. Basically, the cabin is pressurized not by letting more air in, but preventing air from leaving.

Also, there is no separate "AC" system on a transport. Bleed air from the engines (or turbocompressors on the DC-8, 787) is around 250 C, due to the thermodynamic heat of compression. All air coming into the cabin must be air conditioned by the packs/ACM or else the results would be fatal.

When the aircraft descends for landing, the engines are running at near-idle thrust. So there is not much Bleed Air available to be sent through the packs into the Cabin.
So the effect in the Cabin can be similar to "they switched the air-con off".


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How to estimate the altitude of an aircraft at a given distance from a runway?
What altitude should a heavy class commercial freighter be at on a descent path 17 nautical miles from a runway? (Say SDF airport)

Which miles are we talking about? Nautical miles (used in aviation)? Statute miles? International miles?

International miles.

The final descent phase is flown at approximately three degrees flight path angle.
$17 \cdot \tan(3^\circ) \approx 0.89$
If the aircraft has 17 track miles to go to the runway, and it is on a continuous descent of 3 degrees, it will have an altitude of approximately 0.89 miles above the runway elevation. Assuming nautical miles here, that translates to approximately 5400 feet.
However, it may be that the approach contains horizontal flight segments, in which case the altitude may be lower at 17 track miles to go.
If one is 17 miles from the airport and the aircraft is not flying straight to the runway but has to make one or more turns, then the number of track miles to go will be higher than 17 and consequently the aircraft will be higher.
Instead of estimating, one can use a flight track website such as FlightRadar 24 to get the exact height of the aircraft. Please note that the altitude given on tracking websites is typically derived by a pressure altimeter, uncorrected for the local atmosphere. This pressure altitude is different than the geometric altitude of the aircraft. To acquire the exact (GPS measured) height, a subscription is required.

Straight in, on glidepath, about 6000' above the runway. Straight-in & fast, below glidepath at 3000 to 4000 feet above the runway.
That's based on about 3 miles per thousand feet is a standard glidepath, and it takes in the ballpark of 5 miles at idle power & level flight to slow from 250 knots to a configuration for starting down the glideslope.
Those are ballpark numbers; a heavy aircraft slows less rapidly than a lightweight one does, headwinds can shorten those distances, while a tailwind at altitude can extend them. And there are techniques to slow a little faster than that, and you can leave the power in & slow down over a (much) longer distance if desired.
But for ballpark values, those are good approximations.

And crosswinds can impact that somewhat also.

Nice answer to a somewhat broad question :)

Thanks guys, sorry it was such a broad question.

It all depends...
But in many cases, airplanes adhere to specific patterns - and these can differ from airport to airport. While the normal glideslope is around 3° it might vary.
For a start, one might look up the charts of the airport in question.
But at this distance, most planes are not on the glideslope yet - often they capture is from below at somewhat above 10 nm.
But they might also be on a continuous descent - and it is hard to tell by eye at such a distance if the airplane is flying level or descending.
Crosswind and weight do actually have little impact on the glideslope, as in most cases (at least commercial traffic) this is fixed and in most cases, pilots don't deviate from regulations...

Checking the charts for KSDF ILS RW17L (just the one I picked by random) @16.9 NM the altitude would be at or (slightly) above 5000ft MSL when using the default approach.
See https://aeronav.faa.gov/d-tpp/1907/00239IL17L.PDF


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What do these temperature data listed in the EASA Type Certificate Data Sheet mean?
Could you tell me what is the Maximum Permissible Indicated Turbine Temperature (ITT) and the ambient limit temperature for sea level static thrust?

The Maximum Permissible Indicated Turbine Temperature (ITT) is the maximal operating temperature that you can accept without reducing power or shutting down the engine. This temperature is taken at the turbine entry and the limit is there to ensure good cycle endurance of the turbine blade. Higher temperatures and you will have increased fatigue degradation due to reduced mechanical properties. Now your question is why are there 3 different ones:
The maximum continuous is the maximal operating temperature of the engine for infinite life cycle endurance. At that temperature the engine turbine should not wear for the whole life of the engine. But in certain condition it might happen that the temperature exceed this value and in specific cases it was taken into account during the design.
The take-off temperature is what is going to dictate your power input, obviously you want 100%, but high AOA, slow relative speed means that your engine is not max performing and as you are applying full power, you might end up with a richer air-fuel ration than normal condition and see a temperature above your maximum continuous quite often. But the 5 first minutes of the flight at takeoff are critical so reducing power would be not good for safety, thus you allow for slightly higher temps, for a short amount of time.
And for the Maximum starting temp is the absolute limit you should never exceed. During startup the airflow inside the combustion chamber is not perfectly installed and it might happen that the combustion flame come closer to the turbine in addition to the richer air-fuel ratio than in normal condition, thus seeing increased temperature. But this should not extend for more than a few seconds or you'll have to shut down the engine.
Concerning the sea level static thrust this is the maximum output you can ask from the engine and only up to the max temperature indicated. Above that temperature you would see a decrease in performances to prevent exceeding ITT. This would then resume in reduced takeoff performances.

can we deduce the turbine exit temperature from the ITT?

Yes and no. It mostly depends on how much work the turbine extract from the flow to drive the propeller... If you know that you can determine the exit flow properties.

The ambient temperature limit is the highest ambient temperature at which the engine will be able to provide listed Sea Level Static Thrust. If the air temperature is higher, the engine will reduce its thrust to prevent exceeding the maximal permissible temperature in the turbine.
The ambient temperature limit is important in aircraft operation, because at temperatures above this limit there will be less thrust available and thus the take-off performance is reduced.
The Maximum Permissible Indicated Turbine Temperature (ITT) is the maximum indicated turbine temperature that is allowed. Pilots should monitor the ITT and make sure the maximum permissible is not exceeded.


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How can I use a window for an emergency exit?
A survivor of US-Bangla Airlines Flight 211 crash reports exiting out of a broken window.
My question is, in case you need to break an airplane window, what is the best way to do that?

Perhaps the best way to escape out a window is to be in the cockpit.

On Quora: Is it possible for a human to punch through an airplane window with their bare fist?

I personally would doubt the veracity of the quoted report, since they say

I have no recollection after I got out of the plane

I would wager that their recollection of the moments just prior could be faulty.
Nevertheless, to try address your question, my answer is: you don't need to, and you better not try, you would be wasting time.
Aircraft are designed such that complete evacuation must be completed within 90 seconds using only half of the available doors. Trying to open a breach through a window would take much longer than that, would cause otherwise avoidable injuries, and in general would create an opening that is generally quite small for a person to go through.
Your time and efforts are much better spent in reaching a normal emergency exit.
The only situation where exiting through other passages could have sense, is when the impact has created breaches (large enough to pass through) in the hull for you, but the likelihood of you being conscious then is quite low.

In fact the Wikipedia report says the aircraft in this case broke into several pieces.

Broken window, Blocked corridor by riped-from-the-floor-seats+corpses and an anwful lot of lub are necessary here


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Can a single engine on an A320 provide wing anti ice?
Can a single engine on an A320 provide wing anti ice for both wings in the air ?

I believe there is an engine bleed cross-over valve that can be opened to supply the other side, but I can't find a diagram of it that shows both sides of the anti-ice system.

It would probably be a single point of failure if one engine wouldn't be enough to deice both wings.
For this reason there is a cross bleed selector so that the remaining engine can feed the wing on the opposing side.
Here is a link to a computer based training video of the A320 explaining that a single engine bleed is enough for both wings but not for both air conditioning packs and wing anti ice at the same time. So to use have wing anti ice with a single engine you need to turn off one pack and enable cross bleed.

What is a "pack" ?

@zymhan PACKs are _Pneumatic Air Cycle Kit_s  which are a central part of the A/C system in aircraft cabins and cockpits.

You cannot use APU bleed air on the 320, it is unregulated and too hot for anti-ice. See: Why can't we use APU bleed-air for wing anti-ice on the A320?